The invention relates to a combustion chamber of a gas turbine engine with an upstream fairing for separating the gas stream, to an annular wall forming a cap of the upstream fairing of the chamber, and to a gas turbine engine with the chamber.
A turbojet comprises, from upstream to downstream in the direction of gas flow, a fan, one or more compressor stages, a combustion chamber, one or more turbine stages and a gas exhaust nozzle. The terms “external” and “internal” are intended to mean radially external and internal with respect to the axis of the turbojet. The terms “outer” and “inner” are intended to mean the outer side and the inner side of the combustion chamber.
With reference to FIG. 1, which represents a combustion chamber 1 of the prior art, the combustion chamber 1 is generally annular around the axis of the turbojet. It comprises, in its upstream portion, a chamber end section 2 with injection systems supplied with fuel by injectors 3 connected to a supply line 4. The injection systems are distributed along the chamber end section 2. The gas of the primary stream emerges upstream of the chamber 1 via a diffuser 5, from which the gas stream is separated into a stream 6 passing into the combustion chamber 1 to allow combustion of the fuel injected by the injector 3, referred to as combustion stream 6, into an external bypass stream 7 which externally bypasses the inlet of the chamber 1, and into an internal bypass stream 8 which internally bypasses the inlet of the chamber 1. The streams 7, 8 which bypass the inlet of the chamber are used for cooling the chamber 1, in particular.
The primary gas stream is separated at a fairing 9. This fairing 9 comprises two parts, called an external cap 10 and an internal cap 11. The external cap 10 takes the form of an annular metal sheet domed toward the upstream side, fastened to the combustion chamber 1 at an outer downstream surface portion 15, and the inner upstream edge 12 of which forms a fold in the downstream direction, thus forming an aerodynamic surface for separation into an external bypass stream 7 and a combustion stream 6. Likewise, the internal cap 11 takes the form of an annular metal sheet domed toward the upstream side, fastened to the combustion chamber 1 at an outer downstream surface portion 16, and the inner upstream edge 13 of which forms a fold in the downstream direction, forming an aerodynamic surface for separating the internal bypass stream 8 and the combustion stream 6.
The external 10 and internal 11 caps are fastened on the outer side of the external 31 and internal 32 wall, respectively, of the combustion chamber 1, at their outer downstream surface portion 15, 16, respectively, by bolts 14. The external 10 and internal 11 caps are therefore mounted in cantilever fashion on the combustion chamber 1.
The combustion chamber is subjected to vibrational stresses, particularly as a result of the combustion and the engine speed. The caps 10, 11 are therefore subjected to these vibrations, in particular the external cap 10. The caps 10, 11 are also subjected to other dynamic excitation frequencies, in particular certain harmonic frequencies of the rotational speed of the rotating elements of the turbojet. The caps 10, 11, mounted in cantilever fashion, may have resonance modes close to the aforementioned frequencies and are therefore subjected to high mechanical dynamic stresses. The caps 10, 11 are consequently exposed to the risks of breaking or cracking.
Various solutions to this problem have been proposed.
A first solution involves providing an annular damping ring, housed in the fold 12, 13 of the caps 10, 11 (or only in the fold of the external cap which is most subjected to the vibrational stresses); the fold 12, 13 is to this end wrapped around the ring so as to hold it in place. The friction caused by the presence of the ring provides an effect of damping and therefore of shifting the frequencies of the resonance modes of the caps 10, 11, which enables them to be distanced from the vibrational frequencies to which the caps 10, 11 are subjected. However, such a device has the disadvantage of low mechanical strength. There is a risk of the ring loosening, or even breaking (on account of the vibrations to which it is subjected), which diminishes or cancels out its effectiveness.
A second solution involves providing an integrated fairing 9, which will thus be termed a covering. The external 10 and internal 11 caps are then formed in a single piece, with connection tabs between them at their inner upstream edges 12, 13. Such a device has two disadvantages. First, given the connecting tabs between the caps 10, 11, the flow cross section for the combustion stream 6 is reduced; now it is an established fact that this cross section must be as large as possible so as to promote the flow of this stream in order to achieve better combustion efficiency. Second, it is appropriate for the cutouts between the tabs to be formed by laser cutting, these cutouts having to have the equivalent of the folds 12, 13 around their contour. Producing such an integrated covering is very difficult and therefore expensive.